CN106884687B

CN106884687B - System and method for cooling turbine shroud trailing edges

Info

Publication number: CN106884687B
Application number: CN201611167260.3A
Authority: CN
Inventors: M.L.本杰明; B.P.莱西; D.帕尔; S.J.阮
Original assignee: General Electric Co
Current assignee: General Electric Co PLC
Priority date: 2015-12-16
Filing date: 2016-12-16
Publication date: 2021-07-06
Anticipated expiration: 2036-12-16
Also published as: JP2017110648A; EP3181826A1; JP6929049B2; US10309252B2; CN106884687A; US20170175573A1; DK3181826T3; EP3181826B1

Abstract

A shroud segment (40) includes a body (42), the body (42) including a leading edge (44), a trailing edge (46), first and second side edges (48, 50), and a pair of opposing lateral sides (52, 54). The first lateral side (52) is configured to interface with a cavity (56) having a cooling fluid, while the second lateral side (54) is oriented toward the hot gas flow path (47). The shroud segment (40) includes at least one channel disposed within the body (42), wherein the at least one channel includes: a first portion extending in a first direction from the leading edge (44) to the trailing edge (46) from upstream of the trailing edge (46) toward the trailing edge (46); a second portion extending from the trailing edge (46) to upstream (44) of the trailing edge in a second direction from the trailing edge (46) to the leading edge (44); and a third portion extending in the first direction from upstream of the trailing edge (46) toward the trailing edge (46).

Description

System and method for cooling turbine shroud trailing edges

Technical Field

The subject matter disclosed herein relates to gas turbine engines, and more particularly to turbine shrouds for gas turbine engines.

Background

Turbomachines, such as gas turbine engines, may include a compressor, a combustor, and a turbine. The gas is compressed in a compressor, combined with fuel, and then fed into a combustor where the gas/fuel mixture is combusted. The high temperature and high energy exhaust fluid is then fed along a hot gas path to a turbine where the energy of the fluid is converted into mechanical energy. High temperatures along the hot gas path may heat turbine components (e.g., turbine shrouds), causing degradation of the components.

Disclosure of Invention

Certain embodiments commensurate in scope with the originally presented subject matter are summarized below. These embodiments are not intended to limit the scope of the claimed subject matter, but rather, they are intended to provide a brief summary of possible forms of the subject matter. Indeed, the subject matter may encompass a variety of forms that may be similar to or different from the embodiments described below.

According to a first embodiment, a shroud segment for use in a turbine section of a gas turbine engine is provided. The shroud segment includes a body including a leading edge, a trailing edge, a first side edge, a second side edge, and a pair of opposing lateral sides between the leading edge and the trailing edge and between the first side edge and the second side edge. A first lateral side of the pair of opposing lateral sides is configured to interface with a cavity having a cooling fluid, and a second lateral side of the pair of opposing lateral sides is oriented toward the hot gas flow path. The shroud further includes at least one channel disposed within the body on a second lateral side proximate the trailing edge, wherein the at least one channel includes: a first portion extending in a first direction from the leading edge to the trailing edge from upstream of the trailing edge toward the trailing edge; a second portion extending in a second direction from the trailing edge to the leading edge upstream from the trailing edge; and a third portion extending in the first direction from upstream of the trailing edge toward the trailing edge. At least one channel is configured to receive cooling fluid from the cavity to cool the trailing edge.

According to a second embodiment, a gas turbine engine is provided. The gas turbine engine includes a compressor, a combustion system, and a turbine section. The turbine section includes: a housing; an outer shroud segment coupled to the housing; an inner shroud segment coupled to the outer shroud segment to form a cavity configured to receive cooling fluid from the compressor. The inner shroud segment includes: a body having a front edge, a rear edge, a first side edge, a second side edge, and a pair of opposing lateral sides between the front edge and the rear edge and between the first side edge and the second side edge. A first lateral side of the pair of opposing lateral sides is configured to interface with the cavity and a second lateral side of the pair of opposing lateral sides is oriented toward the hot gas flow path. The inner shroud segment includes a plurality of channels disposed within the body on a second lateral side proximate the trailing edge, wherein each channel is arranged in a serpentine pattern. The plurality of channels are configured to receive cooling fluid from the cavity to cool the trailing edge.

According to a third embodiment, a shroud segment for use in a turbine section of a gas turbine engine is provided. The shroud segment includes a body including a leading edge, a trailing edge, a first side edge, a second side edge, and a pair of opposing lateral sides between the leading edge and the trailing edge and between the first side edge and the second side edge. A first lateral side of the pair of opposing lateral sides is configured to interface with a cavity having a cooling fluid, and a second lateral side of the pair of opposing lateral sides is oriented toward the hot gas flow path. The shroud segment also includes a plurality of channels disposed within the body on a second lateral side adjacent the trailing edge, wherein each channel is arranged in a serpentine pattern and each channel includes a free end disposed upstream of the trailing edge. The shroud segment also includes a plurality of inlet passages. A respective inlet passage of the plurality of inlet passages is coupled to a respective free end of a respective channel of the plurality of channels upstream of the trailing edge, wherein the respective inlet passage extends from the respective free end to the first lateral side, and the respective inlet passage is configured to provide cooling fluid from the cavity to the respective channel to cool the trailing edge.

A first aspect of the present invention provides a shroud segment for use in a turbine section of a gas turbine engine, comprising: a body comprising a leading edge, a trailing edge, a first lateral edge, a second lateral edge, and a pair of opposing lateral sides between the leading edge and the trailing edge and between the first lateral edge and the second lateral edge, wherein a first lateral side of the pair of opposing lateral sides is configured to interface with a cavity having a cooling fluid and a second lateral side of the pair of opposing lateral sides is configured to be oriented toward a hot gas flow path; and at least one channel disposed within the body on the second lateral side proximate the trailing edge, wherein the at least one channel comprises: a first portion extending in a first direction from the leading edge to the trailing edge from upstream of the trailing edge toward the trailing edge; a second portion extending in a second direction from the trailing edge to the leading edge upstream from the trailing edge; and a third portion extending in the first direction from upstream of the trailing edge toward the trailing edge; and wherein the at least one passage is configured to receive the cooling fluid from the cavity to cool the trailing edge.

A second aspect of the present invention is the first aspect wherein the first portion, the second portion, and the third portion are linear.

A third aspect of the present invention is the second aspect wherein the at least one passage includes a first curved portion coupling the first portion to the second portion near the trailing edge.

A fourth aspect of the present invention is the third aspect wherein the at least one passage includes a second curved portion coupling the second portion to the third portion upstream of the trailing edge.

A fifth aspect of the present invention is the first aspect wherein the first portion, the second portion, and the third portion are parallel with respect to each other.

A sixth aspect of the present invention is the first aspect wherein the at least one channel is arranged in a serpentine pattern.

A seventh aspect of the present invention is the first aspect wherein the first portion includes a first free end, the shroud segment includes an inlet passage coupled to the first free end and extending in a radial direction from the first free end to the first lateral side, and the inlet passage is configured to provide the cooling fluid from the cavity to the at least one channel.

An eighth aspect of the present invention is the seventh aspect wherein the third portion includes a second free end disposed at the trailing edge, and the at least one passage is configured to discharge the cooling fluid from the body at the trailing edge via the second free end.

A ninth aspect of the present invention is the first aspect wherein the first portion is located closer to a central axis of the body extending from the leading edge to the trailing edge than the second portion and the third portion.

A tenth technical means is the ninth technical means, wherein the second portion is located closer to the central axis than the third portion.

An eleventh aspect of the present invention is the first aspect wherein the at least one passage is electro-discharge machined into the body.

A twelfth aspect of the present invention is the first aspect wherein the body has a length from the leading edge to the trailing edge, and at least the first passages are all disposed within a last quarter of the length.

A thirteenth aspect of the present invention is in the first aspect, comprising a pre-sinter preform layer brazed to the second lateral side, wherein the pre-sinter preform layer comprises a first surface configured to interface with the hot gas flow path and a second surface defining the at least one channel with the body.

A fourteenth aspect of the present invention provides a gas turbine engine, comprising: a compressor; a combustion system; and a turbine section comprising: a housing; an outer shroud segment coupled to the outer shell; an inner shroud segment coupled to the outer shroud segment to form a cavity configured to receive cooling fluid from the compressor, wherein the inner shroud segment comprises: a body comprising a leading edge, a trailing edge, a first lateral edge, a second lateral edge, and a pair of opposing lateral sides between the leading edge and the trailing edge and between the first lateral edge and the second lateral edge, wherein a first lateral side of the pair of opposing lateral sides is configured to interface with the cavity and a second lateral side of the pair of opposing lateral sides is oriented toward a hot gas flow path; a plurality of channels disposed within the body on the second lateral side proximate the trailing edge, wherein each channel is arranged in a serpentine pattern; and wherein the plurality of channels are configured to receive the cooling fluid from the cavity to cool the trailing edge.

A fifteenth aspect of the present invention is the fourteenth aspect, wherein each of the plurality of channels includes a first free end and a second free end, the inner shroud segment includes a respective inlet passage of a plurality of inlet passages coupled to the respective first free ends of the respective channels of the plurality of channels, the plurality of inlet passages extend from the respective first free ends to the first lateral side, and the plurality of inlet passages are configured to provide the cooling fluid from the cavity to the plurality of channels.

A sixteenth technical means is the fifteenth technical means, wherein each of the second free ends of the plurality of channels is disposed at the trailing edge, and the plurality of channels is configured to discharge the cooling fluid from the body at the trailing edge via the second free ends.

A seventeenth technical means is the fourteenth technical means, wherein the body has a length from the leading edge to the trailing edge, and the plurality of channels are all disposed within a last quarter of the length.

An eighteenth aspect of the present invention is the fourteenth aspect, comprising a pre-sinter preform layer brazed to the second lateral side, wherein the pre-sinter preform layer comprises a first surface configured to interface with the hot gas flow path and a second surface defining the plurality of channels with the body.

A nineteenth aspect of the present invention provides a shroud segment for use in a turbine section of a gas turbine engine, comprising: a body comprising a leading edge, a trailing edge, a first lateral edge, a second lateral edge, and a pair of opposing lateral sides between the leading edge and the trailing edge and between the first lateral edge and the second lateral edge, wherein a first lateral side of the pair of opposing lateral sides is configured to interface with a cavity having a cooling fluid and a second lateral side of the pair of opposing lateral sides is oriented toward a hot gas flow path; a plurality of channels disposed within the body on the second lateral side proximate the trailing edge, wherein each channel is arranged in a serpentine pattern and each channel includes a free end disposed upstream of the trailing edge; and a plurality of inlet passages, wherein respective inlet passages of the plurality of inlet passages are coupled to respective free ends of respective channels of the plurality of channels upstream of the trailing edge, wherein the respective inlet passages extend from the respective free ends to the first lateral side, and the respective inlet passages are configured to provide the cooling fluid from the cavity to the respective channels to cool the trailing edge.

A twentieth aspect of the present invention is the nineteenth aspect wherein the body has a length from the leading edge to the trailing edge, and the plurality of channels are all disposed within a last quarter of the length.

Drawings

These and other features, aspects, and advantages of the present subject matter will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:

FIG. 1 is a block diagram of an embodiment of a turbine system having a turbine shroud with cooling channels;

FIG. 2 is a perspective view of an embodiment of an inner turbine shroud segment coupled to an outer turbine shroud segment;

FIG. 3 is a bottom view (e.g., a view of a lateral side oriented toward a hot gas flow path) of an embodiment of an inner turbine shroud segment;

FIG. 4 is a top view of an embodiment of an inner turbine shroud segment (e.g., a view of a lateral side interfacing with a cavity);

FIG. 5 is a bottom view of an embodiment of an inner turbine shroud segment having a zigzag arrangement of cooling channels near the trailing edge (e.g., a view of the lateral side oriented toward the hot gas flow path);

FIG. 6 is a bottom view (e.g., a view of a lateral side oriented toward a hot gas flow path) of an embodiment of an inner turbine shroud segment having a serpentine arrangement of cooling channels near a trailing edge;

FIG. 7 is a perspective cross-sectional view of an embodiment of a portion of the inner turbine shroud segment of FIG. 5 taken along line 7-7 (with the inlet passages and channels shown in phantom); and

FIG. 8 is a flow diagram of an embodiment of a method for manufacturing an inner turbine shroud segment.

Parts list

10 turbine system

12 fuel nozzle

14 supply of fuel

16 burner

18 turbine

19 turbine shell

20 exhaust outlet

22 shaft

24 compressor

26 air inlet

28 load

30 axial direction

32 radial direction

34 circumferential direction of the rotor

36 inner turbine shroud segment

38 outer turbine shroud segment

40 turbine shroud segment

42 body

44 leading edge

46 trailing edge

47 hot gas flow path

48 first side edge

50 second side edge

52 lateral side

54 lateral side

56 chamber

58 pre-sintered preform layer

60 first surface

62 second surface

74 channel

76 first end portion

78 hook-shaped part

80 free end

82 second end portion

84 arrow head

86 channel

88 channel

90 length

92 opening

93 opening

96 sectional type passageway

98 bridge part

100 part (C)

102 part (c)

104 path

105 outlet hole

106 channels

108 central axis

110 first free end

112 second free end

114 arrow head

116 innermost passageway

118 channel

120 channel

122 channel

124 channel

126 channel

128 channels

130 first part

132 second part

134 third part

136 first curved portion

138 second curved portion

140 first part

142 second part

144 curved portion

146 channel

148 first side

150 second side

152 first free end

154 second free end

156 arrow head

158 first part

160 second part

162 third part

164 first curved portion

166 second curved portion

168 inlet passage

170 outlet orifice

172 method

174 step

Step 176

178 step

180 step

182 step

184 step

186 step

188 step

190 step

192 step

194 step

196 step

Step 198.

Detailed Description

One or more specific embodiments of the present subject matter will be described below. In an effort to provide a concise description of these embodiments, all features of an actual implementation may not be described in the specification. It should be appreciated that in the development of any such actual implementation, as in any engineering project, numerous implementation-specific decisions must be made to achieve the developers' specific goals, such as compliance with system-related and business-related constraints, which may vary from one implementation to another. Moreover, it should be appreciated that such a development effort might be complex and time consuming, but would nevertheless be a routine undertaking of manufacture and manufacture for those of ordinary skill having the benefit of this disclosure.

When introducing elements of various embodiments of the present subject matter, the articles "a," "an," "the," and "said" are intended to mean that there are one or more of the elements. The terms "comprising," "including," and "having" are intended to be inclusive and mean that there may be additional elements other than the listed elements.

The present disclosure is directed to systems and methods for cooling components of a turbine (e.g., a turbine shroud) disposed along a hot gas flow path. Specifically, the inner turbine shroud segment includes a body including a near-surface channel (e.g., a microchannel) disposed on a lateral side oriented toward the hot gas flow path. In certain embodiments, the channel may be disposed near a trailing edge of the body. The pre-sintered preform layer disposed (e.g., brazed) on the lateral side defines a channel with the body. Each channel includes: a first portion extending in a first direction from the leading edge to the trailing edge of the body from upstream of the trailing edge toward the trailing edge; a second portion extending in a second direction from the trailing edge to the leading edge upstream from the trailing edge; and a third portion extending in the first direction from upstream of the trailing edge toward the trailing edge. In certain embodiments, the first, second, and third portions are coupled via a curved portion. In certain embodiments, the individual channels near the trailing edge may be arranged in a serpentine pattern. The channel near the trailing edge is configured to receive cooling fluid (e.g., discharge air from a compressor or impingement air) from a cavity (e.g., a bathtub) defined by the inner turbine shroud segment and an outer turbine shroud segment coupled to the inner turbine shroud segment via an inlet passage coupled to a respective free end of a first portion of the channel extending to a lateral side of the inner turbine shroud segment interfacing with the cavity (i.e., a lateral side opposite the lateral side oriented toward the hot gas flow path). The channels discharge cooling fluid (e.g., spent cooling fluid) from the trailing edge of the body via respective free ends of the third portion of the channels. The shape of the channels provides a larger cooling area near the trailing edge (e.g., larger than typical cooling systems for turbine shrouds) while maintaining flow at a minimum. The shape of the channels is also optimized to provide sufficient cooling in the event of blockage of the channels. The disclosed embodiments of the inner turbine shroud segment may allow the inner turbine shroud segment to be cooled with less air (e.g., less than typical cooling systems for turbine shrouds), resulting in a cost reduction associated with the inflatable air for cooling.

Turning to the drawings, FIG. 1 is a block diagram of an embodiment of a turbine system 10. As described in detail below, the disclosed turbine system 10 (e.g., a gas turbine engine) may use a turbine shroud having cooling channels, described below, that may reduce stress patterns of hot gas path components and improve the efficiency of the turbine system 10. The turbine system 10 may use liquid or gas fuels (such as natural gas and/or hydrogen-rich syngas) to drive the turbine system 10. As shown, the fuel nozzles 12 intake a fuel supply 14, mixing the fuel with an oxidant, such as air, oxygen-enriched air, oxygen-depleted air, or any combination thereof. Although the following discussion refers to the oxidant as air, any suitable oxidant may be used in conjunction with the disclosed embodiments. Once the fuel and air are mixed, the fuel nozzles 12 distribute the fuel-air mixture into the combustor 16 in a suitable ratio for optimal combustion, emissions, fuel consumption, and power output. The turbine system 10 may include one or more fuel nozzles 12 located within one or more combustors 16. The fuel-air mixture is combusted in a chamber within combustor 16, thereby generating hot pressurized exhaust gases. The combustor 16 channels the exhaust gases (e.g., hot pressurized gases) through a transition piece into turbine nozzles (or "stage one nozzles"), as well as buckets (or blades) and nozzles of other stages, causing the turbine 18 to rotate within a turbine casing 19 (e.g., outer casing). The exhaust gas flows toward the exhaust outlet 20. As the exhaust gases pass through the turbine 18, the gases force turbine buckets (or blades) to rotate a shaft 22 along an axis of the turbine system 10. As shown, the shaft 22 may be connected to various components of the turbine system 10, including a compressor 24. The compressor 24 also includes blades coupled to the shaft 22. As the shaft rotates 22, blades within the compressor 24 also rotate, thereby compressing air from an air intake 26 through the compressor 24 and into the fuel nozzles 12 and/or combustor 16. A portion of the compressed air (e.g., exhaust air) from the compressor 24 may be diverted to the turbine 18 or a component thereof without passing through the combustor 16. The exhaust air (e.g., cooling fluid) may be used to cool turbine components, such as nozzles on the shrouds and stators, as well as buckets, disks, and diaphragms on the rotor. The shaft 22 may also be connected to a load 28, which may be a vehicle or a stationary load, such as a generator in a power plant or a propeller on an aircraft. The load 28 may include any suitable device capable of being powered by the rotational output of the turbine system 10. The turbine system 10 may extend along an axial axis or direction 30, a radial direction 32 toward or away from the axis 30, and a circumferential direction 34 about the axis 30. In embodiments, hot gas components (e.g., turbine shrouds, nozzles, etc.) are located in the turbine 18, wherein the hot gas flows across the components causing creep, oxidation, wear, and thermal fatigue of the turbine components. The turbine 18 may include one or more turbine shroud segments (e.g., inner turbine shroud segments) having cooling passages (e.g., near-surface microchannels) to allow for control of the temperature of hot gas path components (e.g., using less cooling air than typical cooling systems for the shroud), to reduce hazard patterns in the components, to extend the useful life of the components (while performing their intended functions), to reduce costs associated with operating the turbine system 10, and to improve the efficiency of the gas turbine system 10.

FIG. 2 is a perspective view of an implementation of an inner turbine shroud segment 36 coupled to an outer turbine shroud segment 38 to form a turbine shroud segment 40. The turbine 18 includes a plurality of turbine shroud segments 40 that together form a respective ring around a respective turbine stage. In certain embodiments, the turbine 18 may include a plurality of inner turbine shroud segments 36 coupled to respective outer turbine shroud segments 38 of each turbine shroud segment 40 disposed in the circumferential direction 34 about the axis of rotation of the turbine 18 (and turbine stages). In other embodiments, the turbine 18 may include a plurality of inner turbine shroud segments 38 coupled to outer turbine shroud segments 38 to form turbine shroud segments 40.

As shown, the inner turbine shroud segment 40 includes a body 42 having an upstream or leading edge 44 and a downstream or trailing edge 46 that both interface with a hot gas flow path 47. The body 42 also includes a first side edge 48 (e.g., a first bevel) and a second side edge 50 (e.g., a second bevel) disposed opposite the first side edge 48, both extending between the leading edge 44 and the trailing edge 46. The body 42 also includes a pair of opposing

lateral sides

52,54 extending between the leading and trailing

edges

44, 46 and the first and second lateral edges 48, 50. In certain embodiments, the body 42 (and specifically the lateral sides 52,54) may be arcuate in the circumferential direction 34 between the first and second lateral sides 48,50 and/or in the axial direction 30 between the leading and trailing

edges

44, 46. The lateral sides 52 are configured to interface with a cavity 56 defined between the inner and outer

turbine shroud segments

36, 38. The lateral sides 54 are configured to be oriented toward the hot gas flow path 47 within the turbine 18.

As described in greater detail below, the body 42 may include a plurality of channels (e.g., cooling channels or microchannels) disposed within the lateral side 54 to facilitate cooling of hot gas flow path components (e.g., the turbine shroud 40, the inner turbine shroud segment 36, etc.). In certain embodiments, some of these channels are disposed near trailing edge 46 with or without other channels disposed in the lateral sides on other portions of body 42. A pre-sintered preform (PSP) layer 58 may be disposed (e.g., brazed) on the lateral side 54 such that a first surface 60 of the PSP layer 58 defines (e.g., surrounds) the channel with the body 42 and a second surface 62 of the PSP layer 58 interfaces with the hot gas flow path 47. The PSP layer 58 may be formed of a superalloy and a braze material. In some embodiments, as an alternative to the PSP layer 58, a non-PSP metal sheet may be provided on the lateral side 54, which defines the channel with the body 42. In certain embodiments, the channels may be cast entirely within the body 42 near the lateral sides 54. In certain embodiments, as an alternative to the PSP layer 58, a barrier coating or thermal barrier coating bridge may be used to surround the passage within the body 42.

In certain embodiments, the body 42 includes a hook portion to allow the inner turbine shroud turbine segment 36 to be coupled to the outer turbine shroud segment 38. As described above, the lateral sides 52 of the inner turbine shroud segment 36 and the outer turbine shroud segment 38 define the cavity 56. The outer turbine shroud segment 38 is generally adjacent to relatively cool fluid or air in the turbine 18 from the compressor 24 (i.e., cooler than the temperature in the hot gas flow path 47). The outer turbine shroud segment 38 includes a passage (not shown) to receive cooling fluid or air from the compressor 24, the compressor 24 providing the cooling fluid to the cavity 56. As described in more detail below, the cooling fluid flows to the channels within the body 42 of the inner turbine shroud segment 36 via inlet passages provided in the body 42 that extend from the lateral sides 52 to the channels. Each channel (disposed in a region not near the trailing edge) includes: a first end portion including a hook portion having a free end, and a second end portion. The second end may include a metering feature (e.g., a portion of the body 42 extending into the channel that narrows the cross-sectional area of a portion of the channel relative to an adjacent cross-sectional area of the channel) to regulate the flow of cooling fluid within the channel. In certain embodiments, each channel itself (excluding the second end) serves as a metering feature (e.g., including a portion of the body 42 extending into the channel). In other embodiments, the inlet passage coupled to the hook portion may include a metering feature (e.g., a portion of the body 42 extending into the inlet passage). In certain embodiments, the channel itself, the second end, or the inlet passage, or a combination thereof, includes a metering feature. In addition, the cooling fluid exits the channel (and the body 42) via the second end at the first side edge 48 and/or the second side edge 50. In certain embodiments, the channels may be arranged in a staggered pattern, wherein the channels have first ends disposed adjacent first side edge 48 and second ends disposed adjacent second side edge 50, while adjacent channels have opposite orientations (i.e., a first end disposed adjacent second side edge 50 and a second end disposed adjacent first side edge 48). The hooked portion of the channel provides a larger cooling area (e.g., larger than typical cooling systems for turbine shrouds) by adding a segment of cooling channel near the slashface while keeping flow to a minimum. In addition, the hook-shaped portion allows for better spacing of the straight portions of the channel. In certain embodiments, the body 42 includes a channel disposed near the trailing edge 46 that is shaped differently than a channel disposed on the remainder of the body 42. For example, the channels near trailing edge 46 (which will be described in more detail below) may each include a serpentine pattern. The shape of the channels is also optimized to provide sufficient cooling in the event of blockage of the channels. The disclosed embodiments of the inner turbine shroud segment may allow the inner turbine shroud segment to be cooled with less air (e.g., less than typical cooling systems for turbine shrouds), resulting in a cost reduction associated with inflatable air for cooling.

FIG. 3 is a bottom view of an embodiment of the inner turbine shroud segment 36 without the PSP layer 58 (e.g., a view of the lateral side 54 of the body 42 oriented toward the hot gas flow path). As shown, the body 42 includes a plurality of channels 74 (e.g., cooling channels or micro-channels) disposed within the lateral side 54. The body 42 may include 2 to 40 or more passages 74 (as shown, the body 42 includes 23 passages 74). Each channel 74 is configured to receive cooling fluid from the cavity 56. Each channel 74 includes a first end 76 that includes a hook portion 78 having a free end 80. Each hook-shaped portion 78 has a hook angle radius ranging from about 0.05 to 4 millimeters (mm), 0.1 to 3mm, 1.14 to 2.5mm, and all subranges therebetween. As described in greater detail below, the free end 80 of each hook portion 78 is coupled to an inlet passage that allows the passage 74 to receive cooling fluid from the cavity 56. The curvature of the hook portion 78 allows more channels 74 to be provided in the lateral side 54. In addition, the hooked portion 78 provides a larger cooling area (e.g., larger than typical cooling systems for turbine shrouds) by adding a segment of the cooling passage 74 near the side edges 48,50 while maintaining flow at a minimum. In addition, the hook portion 78 allows for better spacing of the straight portions of the channel 74. Moreover, the swivel of the hook portions 78 allows the straight portions of the channels to be consistently spaced from adjacent channels to hook over the main portion of the body 42 of the shroud segment 36. In certain embodiments, the hook portions 78 may be adjusted to allow the spacing of the straight portions of the channels 74 to be more closely packed for higher heat load zones. In general, the shape of the channels 74 is also optimized to provide adequate cooling in the event of plugging of the channels 74. Each channel 74 also includes a second end 82 that allows the spent cooling fluid to exit the body 42 via the side edges 48,50 via the exit holes as indicated by arrow 84. In certain embodiments, the second end 82 includes a metering feature configured to regulate (e.g., meter) the flow of cooling fluid within the respective passage 74. In certain embodiments, each channel 74 may form a segmented channel at the second end 82. Specifically, a bridge portion of the body 42 may extend across each channel 74 within the second end 82 (e.g., in a direction from the leading edge 44 to the trailing edge 46), with a portion of the channel 74 upstream of the bridge portion and a portion of the channel 74 downstream of the bridge portion. The passages may extend below the bridge portion that fluidly connects the portions of the channels 74 upstream and downstream of the bridge portion. In certain embodiments, each channel 74 itself (excluding the second end 82) serves as a metering feature (e.g., includes a portion of the body 42 extending into the channel). In other embodiments, the inlet passage coupled to the hook portion 78 may include a metering feature (e.g., a portion of the body 42 extending into the inlet passage). In certain embodiments, the channel 74 itself, the second end 82, or the inlet passage, or a combination thereof, includes a metering feature.

As shown, some of the channels 74 (e.g., channel 86) include a hook portion 78 of the first end 76 disposed adjacent the side edge 50 and a second end 82 disposed adjacent the side edge 48, while some of the channels 74 (e.g., channel 88) include a hook portion 78 of the first end 76 disposed adjacent the side edge 48 and a second end 82 disposed adjacent the side edge 50. In certain embodiments, the channels 74 are arranged in a staggered pattern (e.g., channels 86,88), with one channel 74 having a hook portion 78 disposed near one

side edge

48 or 50 and a second end portion 82 disposed near the

opposite side edge

48 or 50, with adjacent channels 74 having an opposing orientation. As shown, the channel 74 extends between the side edges 48,50 from adjacent the leading edge 44 to adjacent the trailing edge 46. In some embodiments, the channel 74 may extend between the side edges 48,50 covering approximately 50 to 90 percent, 50 to 70 percent, 70 to 90 percent, and all subranges therein of the length 90 of the body 42 between the leading edge 44 and the trailing edge 46. For example, the channel 74 may extend between the side edges 48,50 covering approximately 50, 55, 60, 65, 70, 75, 80, 85, or 90 percent of the length 90. This allows for cooling along both side edges 48,50, as well as cooling across a significant portion of the body 42 between the leading edge 44 and the trailing edge 46 and side edges 48,50 (specifically, the lateral sides 54 oriented toward the hot gas flow path 47).

FIG. 4 is a top view of an embodiment of the inner turbine shroud segment 36 (e.g., a view of the lateral side 52 interfacing with the cavity 56). As shown, the body 42 includes a plurality of openings or apertures 92 that allow cooling fluid to flow from the cavity 56 into the channel 74 via the inlet passageway. The body also includes a plurality of openings or apertures 93 that allow cooling fluid to flow from the cavity 56 into a channel (as opposed to the channel 74) disposed near the trailing edge 46. In certain embodiments, the inlet passage extends generally in the radial direction 32 from the free end 80 of the hook portion 78 of the channel 74 to the lateral side 52 to allow the cooling fluid to flow into the channel 74. In certain embodiments, the inlet passage may be angled with respect to the lateral side 54. For example, the angle of the inlet passage may range between approximately 45 and 90 degrees, 45 and 70 degrees, 70 and 90 degrees, and all subranges therein.

FIG. 5 is a bottom view (e.g., view of the lateral side 54 oriented toward the hot gas flow path) of an embodiment of the inner turbine shroud segment 36 (without the PSP layer 58) having a zigzag arrangement of cooling channels 106 near the trailing edge 46. As shown, the body 42 includes a plurality of channels 106 (e.g., cooling channels or microchannels) disposed within the lateral side 54 near the trailing edge 46. The body 42 may include 2 to 30 or more channels 106 (as shown, the body 42 includes 13 channels 106). Each channel 106 is configured to receive cooling fluid from the cavity 56 via a first free end 110 and to discharge spent cooling fluid via an exit aperture at the trailing edge 46 via a second free end 112 as indicated by arrow 114. In certain embodiments, the passage 106 may include metering features as described above with respect to the passage 74. The innermost passage 116 extends along the central axis 108 in the axial direction 30 from upstream of the trailing edge 46 to the trailing edge 46. The channels 106 also include a channel 118 flanking the channel 116, a channel 120 flanking the channel 116 and the channel 118, a channel 122 flanking the channel 116 and the channels 118,120, a channel 124 flanking the channel 116 and the channels 118,120,122, a channel 126 flanking the channel 116 and the channels 118,120,122,124, and a channel 128 flanking the channel 116 and the channel 118,120,122,124,126. The passages 118,120,122,124,126 each include a first portion 130 having a first free end 110 that extends along the axial direction 30, parallel to the central axis 108, from upstream of the trailing edge 46 toward the trailing edge 46. The passages 118,120,122,124, and 126 also each include a second portion 132 that extends perpendicular to the central axis 108 and away from the central axis 108 (as well as parallel to the trailing edge 46). The passages 118,120,122,124, and 126 also each include a third portion 134 having the second free end 112 that extends along the axial direction 30, parallel to the central axis 108, from upstream of the trailing edge 46 toward the trailing edge 46. Channels 118,120,122,124, and 126 each include a first curved portion 136 that couples first portion 130 and second portion 132, and a second curved portion 138 that couples second portion 132 and third portion 134. The portions 130 are parallel with respect to each other. Also, the portions 132 are parallel with respect to each other. Also, the portions 134 are parallel with respect to each other. Each of the passages 128 includes a first portion 140 having the first free end 110 that extends perpendicular to the central axis 108 and away from the central axis 108 (and parallel to the trailing edge 46). The passages 128 also each include a second portion 142 having the second free end 112 that extends along the axial direction 30, parallel to the central axis 108, from upstream of the trailing edge 46 to the trailing edge 46. Each of the channels 128 includes a curved portion 144 that couples the first and

second portions

140, 142. Portion 140 is parallel with respect to portion 132. And portion 142 is parallel with respect to portion 132. The entirety of the channels 106 may be disposed within the last approximately 25 percent of the length 90 of the body 42 near the trailing edge 46. In certain embodiments, the channel 106 may be disposed within the last approximately 15 to 25 percent of the length 90 of the body 42 near the trailing edge 46. The passages 106 provide a larger cooling area near the trailing edge 46 (e.g., larger than typical cooling systems for turbine shrouds) while maintaining flow minimization. In general, the shape of the channels 106 is also optimized to provide sufficient cooling in the event of plugging of the channels 106. In certain embodiments, the body 42 may include only the channel 106 (as opposed to both the channels 74,106).

FIG. 6 is a bottom view (e.g., a view of a lateral side oriented toward the hot gas flow path) of an embodiment of an inner turbine shroud segment having a serpentine arrangement of cooling channels 146 near trailing edge 46. As shown, the body 42 includes a plurality of channels 146 (e.g., cooling channels or microchannels) disposed within the lateral side 54 near the trailing edge 46. The body 42 may include 2 to 30 or more channels 146 (as shown, the body 42 includes 10 channels 146). The channel 146 is disposed about the central axis 108 of the body 42. As shown, 5 of the channels 146 are disposed on a first side 148 of the central axis 108, while the other 5 channels 146 are disposed on a second side 150 of the central axis 108, with their orientations flipped 180 degrees with respect to the channels 146 on the first side 148 (e.g., to form a mirror image about the central axis 108). Each channel 146 is configured to receive cooling fluid from the cavity 56 via the first free end 152 and to discharge spent cooling fluid via the second free end 154 via the exit aperture at the trailing edge 46 as indicated by arrow 156. In certain embodiments, the channel 146 may include metering features as described above with respect to the channel 74. Each passage 146 includes a first portion 158 having the first free end 152 that extends along the axial direction 30, parallel to the central axis 108, from upstream of the trailing edge 46 toward the trailing edge 46. Each passage 146 also includes a second portion 160 that extends from near trailing edge 46 (e.g., parallel to central axis 108) to upstream of the second edge (e.g., opposite direction 30). Each passage 146 also includes a third portion 162 that extends along axial direction 30, parallel to central axis 108, from upstream of trailing edge 46 toward trailing edge 46. The first portion 158, the second portion 160, and the third portion 162 are parallel with respect to one another. As shown, the first portion 158, the second portion 160, and the third portion 162 are linear. The second portion 160 is disposed between the first portion 158 and the third portion 162. The first portion 158 is positioned closer to the central axis 108 than the second portion 160 and the third portion 162. Each channel 146 also includes a first curved portion 164 disposed proximate trailing edge 46 that couples first portion 160 and second portion 162. Each channel 146 also includes a second curved portion 166 disposed upstream of trailing edge 46 that couples second portion 160 and third portion 162. Thus, as shown, each channel 146 includes a serpentine pattern. All of the channels 146 may be disposed within the last approximately 25 percent of the length 90 of the body 42 near the trailing edge 46. In certain embodiments, the channel 146 may be disposed within the last approximately 15 to 25 percent of the length 90 of the body 42 near the trailing edge 46. The channels 146 provide a larger cooling area near the trailing edge 46 (e.g., larger than typical cooling systems for turbine shrouds) while maintaining flow minimization. In general, the shape of the channels 146 is also optimized to provide sufficient cooling in the event of plugging of the channels 146. In certain embodiments, the body 42 may include only the channel 146 (as opposed to both the channels 74,146).

FIG. 7 is a perspective cross-sectional view of an embodiment of a portion of the inner turbine shroud segment 36 of FIG. 5 taken along line 7-7 (with the inlet passage 168 and the channel 106 shown in phantom). As shown, the inlet passageway 168 (shown in phantom) extends generally in the radial direction 32 from the free end 110 of the first portion 130,140 of the channel 106 to the lateral side 52 (e.g., to the opening 93) to allow the cooling fluid to flow into the channel 106. The channel 146 (e.g., the free end 152 of the first portion 158) may also be coupled to an inlet passageway similar to the passageway 168. In certain embodiments, the inlet passageway 168 may be angled with respect to the lateral side 54. For example, the angle of the inlet passageway 168 may range between approximately 45 and 90 degrees, 45 and 70 degrees, 70 and 90 degrees, and all subranges therein. Also, as shown in FIG. 7, the exit holes 170 for the channels 106 (or channels 146) discharge the spent cooling fluid from the trailing edge 46.

FIG. 8 is a flow diagram of an embodiment of a method 172 for manufacturing the inner turbine shroud segment 36. The method 172 includes casting the body 42 (block 174). The method 172 further includes milling the gas path surface onto the body 42 (block 176). Specifically, the lateral side 54 configured to be oriented toward the hot gas flow path 47 may be milled into an arc along the circumferential direction 34 between the first and second side edges 48,50 and/or along the axial direction 30 between the leading and trailing

edges

44, 46. The method 172 also includes forming (e.g., machining, electrodischarge machining, etc.) the channel 74,106,146 to the lateral side 54 of the body 42 (block 178). The method 172 also includes optionally forming (e.g., machining, electric discharge machining, etc.) an outlet feature or outlet flag feature that indicates a location of an outlet hole to be drilled or electric discharge machined in the second end 82 of the channel 74 (or the portion 134,162 of the channel 106,146, respectively) (block 180). The method 172 still further includes forming (e.g., machining, electro-discharge machining, etc.) an access passageway from the lateral side 52 to the free end 80 of the hook portion 78 of the first end 76 of the channel 74 and/or the access passageway 168 to the channel 106,146 (block 182). The method 172 includes covering the openings or

apertures

92,93 of the inlet passage 94,168 (block 184) to block debris from entering the channel 74,106,146 during manufacture of the inner turbine shroud segment 36. The method 172 includes brazing the PSP layer 58 onto the lateral side 54 (block 186) such that the first surface 60 of the PSP layer 58 defines (e.g., surrounds) the channel 74,106,146 with the body 42 and the second surface 62 of the PSP layer 58 interfaces with the hot gas flow path 47. In certain embodiments, as an alternative to the PSP layer 58, a non-PSP metal sheet may be provided on the lateral side 54, which together with the body 42 defines the channels 74,106 and 146. In certain embodiments, as an alternative to the PSP layer 58, a barrier coating or TBC bridge may be used to surround the passage 74,106,146 within the body 42. The method 172 further includes checking brazing of the PSP layer 58 to the body 42 (block 188). The method 172 further includes machining the bevel (e.g., the side edges 48,50) (block 190). The method 172 also includes removing the cover from the

openings

92,93 of the inlet passage 94,168 (block 192). The method 172 even further includes forming (e.g., machining, electro-discharge machining, etc.) the outlet aperture of the second end 82 of the channel 74 to allow the cooling fluid to flow out of the side edges 48,50, and/or the outlet aperture 170 (e.g., out of the metering aperture) (block 194). In certain embodiments, the channel 74,106,146, metering feature, and inlet passage 94 may be cast within the body 42.

Technical effects of the disclosed embodiments include providing a system and method for cooling the trailing edge 46 of the inner turbine shroud segment 36. Specifically, the inner turbine shroud segment 36 includes near-surface microchannels 146 on the lateral side 54 that are enclosed within the body 42 via the PSP layer 58. The channel 146 includes a free end 110 coupled to an inlet passage 168 to allow cooling fluid to flow into the channel 146 to cool the trailing edge 46 of the inner turbine shroud segment 36. The passage 146 may also include metering features to regulate the flow of cooling fluid within the passage 146. The serpentine shape of the channels 146 provides a larger cooling area near the trailing edge 46 (e.g., larger than typical cooling systems for turbine shrouds) while maintaining flow minimization. The shape of the channels 146 is also optimized to provide adequate cooling in the event of plugging of the channels 146. The disclosed embodiments of the inner turbine shroud segment 36 may allow the trailing edge 46 of the inner turbine shroud segment 36 to be cooled with less air (e.g., less than typical cooling systems for turbine shrouds), resulting in a cost reduction associated with inflatable air for cooling.

This written description uses examples to disclose the subject matter, including the best mode, and also to enable any person skilled in the art to practice the subject matter, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the subject matter is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

1. A shroud segment (40) for use in a turbine section (18) of a gas turbine engine (10), comprising:

a body (42) including a leading edge (44), a trailing edge (46), a first side edge (48), a second side edge (50), and a pair of opposing lateral sides (52,54) between the leading edge (44) and the trailing edge (46) and between the first side edge (48) and the second side edge (50), wherein a first lateral side (52) of the pair of opposing lateral sides (52,54) is configured to interface with a cavity (56) having a cooling fluid and a second lateral side (54) of the pair of opposing lateral sides (52,54) is oriented toward a hot gas flow path (47); and

at least one first channel disposed within the body (42) on the second lateral side (54) proximate the trailing edge (46), wherein the at least one first channel comprises: a first portion extending in a first direction from the leading edge (44) to the trailing edge (46) from upstream of the trailing edge (46) toward the trailing edge (46); a second portion extending in a second direction from the trailing edge (46) to the leading edge (44) from the trailing edge (46) to upstream of the trailing edge (46); and a third portion extending in the first direction from upstream of the trailing edge (46) towards the trailing edge (46); and

wherein the at least one first channel is configured to receive the cooling fluid from the cavity (56) to cool the trailing edge (46);

wherein the shroud segment further comprises:

at least one second channel (74) provided in the body (42) on the second lateral side (54) and provided in an area not adjacent to the rear edge (46), wherein the at least one second channel comprises a first end portion (76) and a second end portion, the first end portion (76) comprising a hook-shaped portion (78) having a free end (80) adjacent to one lateral edge (48, 50), the second end portion extending to the other lateral edge (48, 50),

wherein the at least one second passage is coupled to the cavity (56) via an inlet passage coupled to the free end (80) of the hook portion (78) to receive cooling fluid from the cavity (56) to cool the side edges (48, 50).

2. The shroud segment (40) of claim 1, wherein the first portion, the second portion, and the third portion are linear.

3. The shroud segment (40) of claim 2, wherein the at least one first channel includes a first curved portion coupling the first portion to the second portion proximate the trailing edge (46).

4. The shroud segment (40) of claim 3, wherein the at least one first channel includes a second curved portion coupling the second portion to the third portion upstream of the trailing edge (46).

5. The shroud segment (40) of claim 1, wherein the first portion, the second portion, and the third portion are parallel with respect to one another.

6. The shroud segment (40) of claim 1, wherein the at least one first channel is arranged in a serpentine pattern.

7. The shroud segment (40) of claim 1, wherein the first portion includes a first free end, the shroud segment (40) includes an inlet passage coupled to the first free end and extending in a radial direction (32) from the first free end to the first lateral side (52), and the inlet passage is configured to provide the cooling fluid from the cavity (56) to the at least one first channel.

8. The shroud segment (40) of claim 7, wherein the third portion includes a second free end disposed at the trailing edge (46), and the at least one first channel is configured to exhaust the cooling fluid from the body (42) at the trailing edge (46) via the second free end.

9. The shroud segment (40) of claim 1, wherein the first portion is located closer to a central axis of the body (42) extending from the leading edge (44) to the trailing edge (46) than the second portion and the third portion.

10. The shroud segment (40) of claim 9, wherein the second portion is located closer to the central axis than the third portion.

11. The shroud segment (40) of claim 1, wherein the at least one first channel is electro-discharge machined into the body (42).

12. The shroud segment (40) of claim 1, wherein the body (42) has a length from the leading edge (44) to the trailing edge (46), and the at least one first channel is disposed entirely within a last quarter of the length.

13. The shroud segment (40) of claim 1, including a pre-sintered preform layer (58) brazed onto the second lateral side (54), wherein the pre-sintered preform layer (58) includes a first surface configured to interface with the hot gas flow path (47) and a second surface defining the at least one first channel with the body (42).

14. A gas turbine engine (10), comprising:

a compressor (24);

a combustion system (16); and

a turbine section (18) comprising:

a housing (19);

an outer shroud segment (38) coupled to the outer shell (19);

an inner shroud segment (36) coupled to the outer shroud segment (38) to form a cavity (56), the cavity (56) configured to receive a cooling fluid from the compressor (24), wherein the inner shroud segment (36) comprises a shroud segment in accordance with any preceding claim.

15. The gas turbine engine (10) of claim 14, wherein each of the at least one first and at least one second channels includes a first free end and a second free end, the inner shroud segment (36) includes a respective inlet passage of a plurality of inlet passages coupled to the respective first free end of the respective ones of the at least one first and at least one second channels, the plurality of inlet passages extending from the respective first free end to the first lateral side (52), and the plurality of inlet passages are configured to provide the cooling fluid from the cavity (56) to the at least one first and at least one second channels.